Combustor assembly for a turbine engine

ABSTRACT

A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. The liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end. The quench air jets include a plurality of primary stage air jets and a plurality of secondary stage air jets. The plurality of primary stage air jets are each spaced from the plurality of secondary stage air jets along the axial direction and together provide the combustion chamber with a quench airflow.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine engine, ormore particularly to a combustor assembly for a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine general includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gassesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

Traditionally, the combustion section includes a combustor for receivingcompressed air and fuel and combusting the combination to provide theturbine section with the combustion gasses. The fuel and air istypically provided with an equivalence ratio of about 1:1 such thatsubstantially stoichiometric combustion takes place. However, such maylead to relatively high peak temperatures, and further to undesirableamounts of NOx formation.

Accordingly, a combustion section for a gas turbine engine capable ofavoiding these issues would be useful. More specifically, a combustionsection capable of generating combustion gases having a reduced amountof NOx, while efficiently combusting all of the fuel would beparticularly beneficial.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, arich-quench-lean combustor assembly is provided for a gas turbine enginedefining an axial direction and a circumferential direction. Thecombustor assembly includes a fuel nozzle and a dome, the fuel nozzleattached to the dome. The combustor assembly additionally includes aliner attached to or formed integrally with the dome, the liner and thedome together defining at least in part a combustion chamber. The linerextends between a forward end and an aft end. The liner includes aplurality of quench air jets positioned between the forward end and aftend. The quench air jets include a plurality of primary stage air jetsand a plurality of secondary stage air jets. The plurality of primarystage air jets are each spaced from the plurality of secondary stage airjets along the axial direction.

In another exemplary embodiment of the present disclosure, a gas turbineengine is provided. The gas turbine engine defines an axial directionand a circumferential direction. The gas turbine engine includes acompressor section and a turbine section arranged in serial flow order.The gas turbine engine further includes a rich-quench-lean combustorassembly positioned between the compressor section and the turbinesection. The combustor assembly includes a fuel nozzle and a dome, thefuel nozzle attached to the dome. The combustor assembly additionallyincludes a liner attached to or formed integrally with the dome, theliner and the dome together defining at least in part a combustionchamber. The liner extends between a forward end and an aft end. Theliner includes a plurality of quench air jets positioned between theforward end and aft end. The quench air jets include a plurality ofprimary stage air jets and a plurality of secondary stage air jets. Theplurality of primary stage air jets are each spaced from the pluralityof secondary stage air jets along the axial direction.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is a close-up, cross-sectional view of a combustor assembly inaccordance with an exemplary embodiment of the present disclosure.

FIG. 3 is a close-up, cutaway view of a fuel nozzle in accordance withan exemplary embodiment of the present disclosure.

FIG. 4 is a close-up, cross-sectional view of a portion of the exemplarycombustor assembly of FIG. 2.

FIG. 5 is a perspective view of a section of the exemplary combustorassembly of FIG. 2, with a flow sleeve removed for clarity.

FIG. 6 is a close-up, side, cross-sectional view of a primary stage airjet of the exemplary combustor assembly of FIG. 2 in accordance with anexemplary aspect of the present disclosure.

FIG. 7 is a close-up, outside view of the exemplary primary stage airjet of FIG. 6.

FIG. 8 is a forward-looking-aft view of a combustion chamber, dome, andouter liner of the exemplary combustor assembly of FIG. 2, in accordancewith an exemplary aspect of the present disclosure.

FIG. 9 is a forward-looking-aft view of a combustion chamber, dome, andouter liner of a combustor assembly in accordance with another exemplaryembodiment of the present disclosure.

FIG. 10 is a flow diagram of a method for operating a combustor assemblyin accordance with an exemplary aspect of the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms “forward”and “aft” refer to relative positions within a gas turbine engine, withforward referring to a position closer to an engine inlet and aftreferring to a position closer to an engine nozzle or exhaust. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal axis 12 provided for reference), aradial direction R, and a circumferential direction C (i.e., a directionextending about the axial direction A; not depicted). In general, theturbofan 10 includes a fan section 14 and a core turbine engine 16disposed downstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal axis 12 by LP shaft 36 acrossa power gear box 46. The power gear box 46 includes a plurality of gearsfor stepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. For the embodiment depicted, thenacelle 50 is supported relative to the core turbine engine 16 by aplurality of circumferentially-spaced outlet guide vanes 52, and adownstream section 54 of the nacelle 50 extends over an outer portion ofthe core turbine engine 16 so as to define a bypass airflow passage 56therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. Thepressure of the second portion of air 64 is then increased as it isrouted through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration. Additionally, it will be appreciated that inother embodiments, aspects of the present disclosure may be incorporatedinto any other suitable gas turbine engine, such as a suitableaeronautical gas turbine engine (e.g., turboshaft, turboprop, turbojet,etc.), land-based gas turbine engine (e.g., power generation gas turbineengine), aero-derivative gas turbine engine (e.g., marine applications),etc.

Referring now to FIG. 2, a close-up cross-sectional view is provided ofa rich-quench-lean (“RQL”) combustor assembly (“combustor assembly 100”)in accordance with an exemplary embodiment of the present disclosure. Incertain embodiments, for example, the combustor assembly 100 of FIG. 2may be positioned in the combustion section 26 of the exemplary turbofanengine 10 of FIG. 1. Alternatively, however, it may be positioned in anyother suitable gas turbine engine. For example, in other embodiments,the combustor assembly 100 may be incorporated into one or more of aturboshaft engine, a turboprop engine, a turbojet engine, a land-basedgas turbine engine for power generation, an aero-derivative or marinegas turbine engine, etc.

As shown, the combustor assembly 100 generally includes an inner liner102 extending between an aft end 104 and a forward end 106 generallyalong the axial direction A, as well as an outer liner 108 alsoextending between an aft end 110 and a forward end 112 generally alongthe axial direction A. The inner and outer liners 102, 108 are eachattached to or formed integrally with a dome 114. More particularly, forthe embodiment depicted, the inner and outer liners 102, 108 are eachformed integrally with the dome 114, such that the inner liner 102,outer liner 108, and dome 114 form a one-piece combustor liner extendingcontinuously from the aft 104 end of the inner liner 102, to the forwardend 106 of the inner liner 102, across the dome 114, to the forward end110 of the outer liner 108, and to the aft end 112 of the outer liner108. Although not depicted in FIG. 2, the one-piece combustor liner mayadditionally extend along the circumferential direction C. In certainembodiments, the one-piece combustor liner may extend continuously alongthe circumferential direction C, or alternatively, the combustorassembly 100 may include a plurality of one-piece combustor linersarranged along the circumferential direction C. Furthermore, as will bediscussed in greater detail below, the inner and outer liners 102, 108and dome 114 together at least partially define a combustion chamber116, the combustion chamber 116 having a centerline 118 extendingtherethrough.

For the embodiment depicted, the inner liner 102, the outer liner 108,and dome 114 are each formed of a ceramic matrix composite (CMC)material, which is a non-metallic material having high temperaturecapability. Exemplary CMC materials utilized for such liners 102, 108and the dome 114 may include silicon carbide, silicon, silica or aluminamatrix materials and combinations thereof. Ceramic fibers may beembedded within the matrix, such as oxidation stable reinforcing fibersincluding monofilaments like sapphire and silicon carbide (e.g.,Textron's SCS-6), as well as rovings and yarn including silicon carbide(e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and DowCorning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480),and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), andoptionally ceramic particles (e.g., oxides of Si, Al, Zr, Y andcombinations thereof) and inorganic fillers (e.g., pyrophillite,wollastonite, mica, talc, kyanite and montmorillonite). It should beappreciated, however, that in other embodiments, one or more of theinner liner 102, outer liner 108, and dome 114 may be formed of anyother suitable material, such as a traditional metal alloy.

Referring still to FIG. 2, the combustor assembly 100 further includes aflow sleeve 120 enclosing the inner liner 102, outer liner 108, and dome114. The flow sleeve 120 generally includes an outer portion 122, aninner portion 124, and a forward portion 126. For the embodimentdepicted, the combustor assembly 100 is mounted within an outer casingof the gas turbine engine, and more particularly, the flow sleeve 120 ismounted to an outer combustor casing 128 of the gas turbine engine usingone or more mounting features 130. The one or more mounting features130, for the embodiment depicted, are attached to the outer portion 122of the flow sleeve 120. Additionally, although not depicted, thecombustor assembly 100 may include a plurality of such mounting features130 attached to the outer portion 122 of the flow sleeve 120 and spacedalong the circumferential direction C.

For the exemplary flow sleeve 120 depicted, the outer portion 122,forward portion 126, and inner portion 124 are formed integrally as aone-piece flow sleeve, extending continuously. Additionally, althoughnot depicted, the flow sleeve 120 may additionally extend continuouslyalong the circumferential direction C. However, in other embodiments,one or more of the outer portion 122, forward portion 126, and innerportion 124 may be formed separately and attached in any suitablemanner. Additionally, in other embodiments, the flow sleeve 120 may beformed of a plurality of individual flow sleeves 120 spaced along thecircumferential direction C.

As is also depicted in FIG. 2, the combustor assembly 100 is positioneddownstream of a diffuser 132 of the gas turbine engine. The diffuser 132is positioned at an aft end of a combustion section of the gas turbineengine for providing compressed air 134 from the combustion section tothe combustor assembly 100. The combustor assembly 100 further includesa fuel nozzle 135 for receiving a portion of the compressed air 134,mixing the received portion of the compressed air 134 with fuel, andproviding such fuel-air mixture to the combustion chamber 116. As willbe shown and discussed in greater detail below, the fuel nozzle 135depicted is configured as a pre-mix fuel nozzle and the combustorassembly 100 further includes a plurality of fuel nozzles 135 spacedsubstantially evenly along the circumferential direction C (see, e.g.,FIGS. 5, 8). Additionally, it should be appreciated that in certainembodiments, the fuel may be a natural gas (such as methane), LNG,propane, lean methane, high H2 content fuel, a petroleum distillate(such as No. 2 distillate fuel), kerosene, heavy fuel oil, marine dieselfuel, or any other suitable fuel.

Referring now briefly to FIG. 3, a close-up, perspective,cross-sectional view is provided of a pre-mix fuel nozzle 135 inaccordance with one or more embodiments of the present disclosure as maybe incorporated in the exemplary combustor assembly 100 of FIG. 2.

The fuel nozzle 135 of FIG. 3 generally includes a centerbody 136 and anouter sleeve 138 that generally surrounds the centerbody 136. The outersleeve 138 includes one or more radial vanes 140 forming a radialswirler 142. An inner sleeve 144 is disposed between the centerbody 136and the outer sleeve 138. The inner sleeve 144, at least a part of whichis generally disposed upstream of the radial swirler 142, includes acontoured shroud 146 at an aft end. Generally, the contoured shroud 146is aerodynamically contoured to promote mixing of a liquid or gaseousfuel and air. For example, the contoured shroud 146 includes a pluralityof lobes 148.

Moreover, the radial swirler 142 is disposed radially outward of thecontoured shroud 146 and a fuel injection port 150. The fuel injectionport 150 is defined between the inner sleeve 144 and a portion of theouter sleeve 138, and provides for a flow of fuel from a fuel circuit(not depicted) defined by the fuel nozzle 135. As shown, at least onefuel injection port 150 is disposed upstream, or forward, of thecontoured shroud 146.

The positioning of the radial swirler 142 to the contoured shroud 146and fuel injection port 150 is such that compressed air, such as aportion of the compressed air 134 from the diffuser 132 (see FIG. 2),entering through the radial swirler 142 converges and mixes with aliquid or gaseous fuel exiting a fuel injection port 150. The contouredshroud 146 may aid in positioning the fuel exiting the fuel injectionport 150 such that the convergence of air 126 through the radial swirler142 may deliver high levels of fuel-air mixing.

Generally upstream in the fuel nozzle 135 from the contoured shroud 146,an outer surface 152 of the centerbody 136 and an inner surface 154 ofthe inner sleeve 144 includes a plurality axially oriented vanes 156forming an axial swirler 158. The axial swirler 158 may have anygeometry between at least one outer surface 152 of the centerbody 136and at least one inner surface 154 of the inner sleeve 144, and is notlimited to any particular geometry, unless otherwise specified. Neitherthe centerbody 136 nor the inner sleeve 144 is bound to one diametricvalue for its entire structure. Furthermore, the centerbody 136 andsubsequent surrounding features may have other radial cross-sectionalforms, such as an elliptical or polygonal radial cross section.

Notably, a relationship of the outer sleeve 138 to the centerbody 136creates an annular circuit 160 substantially along a length of thecenterbody 136. In the embodiment shown in FIG. 3, the radialcross-sectional area of the annular circuit 160 at a first location isgreater than at a second location downstream of the first location.Thus, the annular circuit 160 may converge as it extends downstreamalong a center axis 162. More specifically, the inner surface 132 of theouter sleeve 138 is converging toward the center axis 162 as the innersurface 132 extends downstream, while the outer surface 152 of thecenterbody 136 is diverging from the center axis 162. In otherembodiments, however, the inner surface 132 of the outer sleeve 138 mayhave any other suitable shape relative to the outer surface 152 of thecenterbody 136.

Referring still to the exemplary fuel nozzle 135 of FIG. 3, thecenterbody 136 includes a first centerbody circuit 164, a secondcenterbody circuit 166, and a third centerbody circuit 168 leading to atleast one circuit outlet 170 to egress a fluid (e.g. liquid or gaseousfuel, air, inert gas, or combination thereof). The second centerbodycircuit 166 and the third centerbody circuit 168 are positionedgenerally co-axial to the first centerbody circuit 164. In anotherembodiment, the second centerbody circuit 166 or the third centerbodycircuit 168 may be tunnels within the centerbody (i.e., not annularcavities), radially outward from a first centerbody circuit 164. Anycombination of centerbody circuits 164, 166, 168 may be fluidlyconnected toward the downstream end of the centerbody 136 beforeegressing through the centerbody outlet 170. In another embodiment, anycenterbody circuit 164, 166, 168, or a combination thereof, may egressindependently to a circuit outlet 170 without fluid interconnection. Theexemplary fuel nozzle 135 may be configured for independent variableflow rates within each centerbody circuit 164, 166, 168. It should beapparent to one skilled in the art that additional centerbody circuits(fourth, fifth . . . Nth) may be installed and arranged in substantiallysimilar manner as the first, second, and third circuits 164, 166, 168described herein.

The exemplary fuel nozzle 135 depicted in FIG. 3, and described herein,may allow for the fuel nozzle 135 to provide a substantially homogenousmixture of fuel and air to the combustion chamber 116 during operationof the combustor assembly 100. As used herein, “substantiallyhomogenous” means more mixed than not, i.e., having a mixedness greaterthan fifty percent (50%). For example, in certain embodiments, the fuelnozzle 135 may be configured to provide a mixture of fuel and air to thecombustion chamber 116 having at least about a seventy percent (70%)mixedness. More specifically, in certain embodiments, the fuel nozzle135 may be configured to provide a mixture of fuel and air to thecombustion chamber 116 having at least about an eighty percent (80%)mixedness. It should be appreciated, that as used herein, terms ofapproximation, such as “about” or “approximately,” refer to being withina ten percent (10%) margin of error.

Moreover, as used herein, the term “mixedness” with respect to a mixtureof fuel and air, refers to a calculation for determining how a fuelspecies varies over a surface. Mixedness may be calculated generallyusing the following formula:

${{Mixedness} = {1 - \frac{\sigma_{f}}{f}}},$where σ_(f) is a standard deviation of mixture fraction and f is a massweighted average of mixture fraction. Each may be calculated in themanner described below.

The mixture fraction, f, is defined in terms of the atomic massfraction, which may be expressed as follows:

${f = \frac{x_{i} - x_{i,{ox}}}{x_{i,{fuel}} - x_{i,{ox}}}},$where x_(i) is the elemental mass fraction for the element, x_(i,ox)denotes the oxidizer, and x_(i,fuel) denotes the value at the fuelstream.

Further, the standard deviation, σ_(f), of the mixture fraction on asurface is computed using the following formula:

${\sigma\; f} = \sqrt{\frac{\sum\limits_{i = 1}^{n}\;\left( {f - f_{0}} \right)^{2}}{n}}$

Further, still, the mass weighted average of the mixture fraction, f, iscalculated by dividing the summation of the value mixture fractionmultiplied by the absolute value of the dot product of the facet areaand momentum vectors by the summation of the absolute value of the dotproduct of the facet area and momentum vectors, as is indicated in thefollowing formula:

$\overset{\_}{f} = {\frac{\sum\limits_{i = 1}^{n}{f_{i}\rho_{i}{{\overset{\rightarrow}{V_{i}} - {\overset{\rightarrow}{A}}_{i}}}}}{\sum\limits_{i = 1}^{n}{\rho_{i}{{\overset{\rightarrow}{V_{i}} - {\overset{\rightarrow}{A}}_{i}}}}}.}$

Mixedness calculated in accordance with the above method may provide amixedness at an outlet of a fuel nozzle.

Further, it should be appreciated, that in other exemplary embodiments,the combustor assembly 100 may have any other suitable fuel nozzle, andthat the present application is not limited to the exemplary fuel nozzle135 depicted in FIG. 3 and described above, unless specifically solimited by the claims.

Referring now back to FIG. 2, as stated, the combustor assembly 100 isconfigured as an RQL combustor assembly 100 (i.e., a combustor assemblyproviding for a rich combustion, a quench air, and subsequently a leancombustion). In order to achieve an initial rich combustion, the fuelnozzle 135 is further configured to provide a mixture of fuel and air tothe combustion chamber 116 having an equivalence ratio of at least 1.5.More specifically, for the embodiment depicted the fuel nozzle 135 isconfigured to provide the combustion chamber 116 with a mixture of fueland air having an equivalence ratio at least about two (2). Notably, asused herein, “equivalence ratio” refers to a ratio of fuel to air.

As depicted, the combustion chamber 116 includes a primary combustionzone 172 and a secondary combustion zone 174. The substantiallyhomogenous mixture of fuel and air, having an equivalence ratio of atleast 1.5, is combusted in the primary combustion zone 172. However,given the relatively high equivalence ratio of such a mixture of fueland air, incomplete combustion occurs (i.e., less than stoichiometriccombustion). In order to complete the combustion process, the combustorassembly 100 is configured to introduce an additional amount of airdownstream of the primary combustion zone 172 forming the secondarycombustion zone 174. More specifically, the inner liner 102 of thecombustor assembly 100 includes a plurality of quench air jets 176positioned between the forward end 106 and the aft end 104, andsimilarly the outer liner 108 of the combustor assembly 100 includes aplurality of quench air jets 176 positioned between the forward 112 endand the aft end 110.

As will be appreciated, a majority of the compressed air 134 from thediffuser 132 is received by the combustor assembly 100 for combustion(i.e., “compressed air for combustion 134A”), while in at least certainembodiments, a portion of the compressed air 134 from the diffuser 132is diverted downstream for cooling operations (not shown). In order toachieve a desired equivalence ratio, along with the desired mixedness ofthe fuel and air mixture provided to the combustion chamber 116, amajority of the compressed air for combustion 134A is introduced intothe combustion chamber 116 through the plurality of quench air jets 176of the inner liner 102 and outer liner 108. More specifically, in atleast certain embodiments, at least about sixty percent (60%) of thecompressed air for combustion 134A is introduced to the combustionchamber 116 through the quench air jets 176 of the inner liner 102 andthe outer liner 108 is a quench airflow 134B. For example, in certainembodiments, at least about seventy percent (70%) of the compressed airfor combustion 134A is introduced to the combustion chamber 116 throughthe quench air jets 176 of the inner liner 102 and the outer liner 108as quench airflow 134B.

The quench airflow 134B introduced through the quench air jets 176 ofthe inner liner 102 and outer liner 108 may mix with the combustiongases from the primary combustion zone 172 in the secondary combustionzone 174. The mixture of quench airflow 134B through the quench air jets176 and combustion gases from the primary combustion zone 172 may resultin a lean combustion mixture within the secondary combustion zone 174.For example, such a mixture of fuel and air may have an equivalenceratio of less than about 0.75. More specifically, in certain embodimentsat least, the secondary combustion zone 174 may define an equivalenceratio of less than about 0.65.

For the embodiment depicted in FIG. 2, the quench air jets 176 of theinner liner 102 and the outer liner 108 are each positionedapproximately halfway along a length of the inner liner 102 and theouter liner 108, respectively (i.e., about halfway between therespective forward ends 106, 112 and aft ends 104, 110). Notably, theexemplary combustion chamber 116 depicted converges aft of the primarycombustion zone 172, and the quench air jets 176 of the inner liner 102and the outer liner 108 are positioned at the convergence. Morespecifically, the exemplary combustor assembly 100 depicted defines aforward height 178 within the combustion chamber 116 between the outerliner 108 and the inner liner 102 at a location forward of the pluralityof quench air jets 176 of the outer liner 108 and the inner liner 102.Additionally, the combustor assembly 100 defines an aft height 180within the combustion chamber 116 between the outer liner 108 and theinner liner 102 at a location aft of the plurality of quench air jets176 of the outer liner 108 and the inner liner 102. A ratio of theforward height 178 to the aft height 180 is at least about 1.75:1. Forexample, in certain embodiments, the ratio of the forward height 178 tothe aft height 180 may be at least about 2:1.

As used herein, the forward height 178 is defined in a direction thatextends perpendicular to the centerline 118 of the combustion chamber116 and intersects with a longitudinal axis of the gas turbine engine(e.g., axis 12 of FIG. 1). Similarly, the aft height 180 is defined in adirection that extends perpendicular to the centerline 118 of thecombustion chamber 116 and intersects with the longitudinal axis of thegas turbine engine. Further, each of the forward height 178 and aftheight 180 refer to the greatest heights in the respective directions,and more particularly, for the embodiment depicted, refer to heightsimmediately forward and immediately aft, respectively, of the quench airjets 176 of the inner liner 102 and outer liner 108.

Briefly, it should also be appreciated, that for the embodimentdepicted, the plurality of quench air jets 176 of the inner liner 102and of the outer liner 108 are each positioned at approximate the sameposition along the centerline 118 of the combustion chamber 116, i.e.,such that they are aligned. However, in other embodiments, the quenchair jets 176 of the inner liner 102 may be offset from the quench airjets 176 of the outer liner 108 along the centerline 118. For example,in certain embodiments, the quench air jets 176 of the inner liner 102may be positioned forward of the quench air jets 176 of the outer liner108, or alternatively, the quench air jets 176 of the inner liner 102may be positioned aft of the quench air jets 176 of the outer liner 108.

Reference will now be made also to FIGS. 4 and 5. FIG. 4 provides aclose-up, cross-sectional view of a section of the exemplary combustorassembly 100 of FIG. 2, and FIG. 5 provides a perspective view of asection of the exemplary combustor assembly 100 of FIG. 2, with the flowsleeve 120 removed for clarity.

As previously discussed, the mixture of fuel and air provided to theprimary combustion zone 172 of the combustion chamber 116 by the fuelnozzles 135 defines a relatively high equivalence ratio. In order tomaintain the desired equivalence ratio, the combustor assembly 100 isconfigured such that substantially no cooling air enters the combustionchamber 116 forward of the quench air jets 176 of the inner liner 102and the outer liner 108. More particularly, for the embodiment depicted,the inner liner 102 defines a forward section 182 extending between thequench air jets 176 and the dome 114 and the outer liner 108 similarlydefines a forward section 184 extending between the quench air jets 176and the dome 114. For the embodiment depicted, the forward section 182of the inner liner 102, the forward section 184 of the outer liner 108,and the dome 114 are each configured to prevent a flow of cooling airfrom entering the combustion chamber 116, and are further configured tobe cooled substantially by one or both of impingement cooling orconvective cooling. More particularly, for the embodiment depicted, eachof the forward section 182 of the inner liner 102, the forward section184 of the outer liner 108, and the dome 114 are free from any coolingholes and are configured to be cooled through a flow of impingement aironto an outer surface 186 of the respective components.

Further, for the embodiment depicted, the impingement air for coolingthe forward section 184 of the outer liner 108, the forward section 182of the inner liner 102, and the dome 114 is provided through impingementcooling holes 188 defined by the flow sleeve 120. The impingementcooling holes 188 through the flow sleeve 120 are all positioned forwardof the quench air jets 176 of the inner and outer liners 102, 108.During operation, a portion of the compressed air for combustion 134Anot provided to the combustion chamber 116 through the fuel nozzles 135flows through the impingement cooling holes 188 defined by the flowsleeve 120, through a chamber 190 (defined between the flow sleeve 120and the outer liner 108, dome 114, and inner liner 102) directly ontothe outer surfaces 186 of the forward section 194 of the outer liner108, the forward section 182 of the inner liner 102, and the dome 114for cooling such components. Notably, while such a cooling method maynot provide a level of cooling attainable through inclusion of coolingholes through the inner and outer liners 102, 108 and/or dome 114, asdiscussed above, the inner and outer liners 102, 108 and dome 114 may beformed of a CMC material. Additionally, as the combustor assembly 100 isconfigured for incomplete combustion in the primary combustion zone, atemperature within the combustion chamber 116 proximate the forwardsections 182, 184 of the inner and outer liners 102, 108 may be reduced.Accordingly, with such an embodiment the components may be capable ofwithstanding the temperatures necessary for operation without inclusionof cooling holes.

Moreover, referring still to FIGS. 4 and 5, for the embodiment depicted,the quench air jets 176 of the outer liner 108 and the quench air jets176 of the inner liner 102 are each configured in a two-stageconfiguration. More particularly, referring first to the quench air jets176 of the outer liner 108, the quench air jets 176 include a pluralityof primary stage air jets 192 and a plurality of secondary stage airjets 194. The plurality of primary stage air jets 192 are each spacedfrom the plurality of secondary stage air jets 194 along the axialdirection A and along the centerline 118.

Further, for the embodiment depicted the plurality of primary stage airjets 192 are each configured as relatively large air jets, while theplurality of secondary stage air jets 194 are each configured asrelatively small air jets. For example, each of the plurality of primarystage air jets 192 define a cross-sectional area 196 (see cross-sectionclose up in FIG. 4) and each of the plurality of secondary stage airjets 194 also defines a cross-sectional area 198 (see cross-sectionclose up in FIG. 5). For the embodiment depicted, the cross-sectionalarea 196 of the primary stage air jets 192 is greater than thecross-sectional area 198 of the secondary stage jets. For example, in atleast certain embodiments, a ratio of the cross-sectional area 196 ofthe primary stage air jets 192 to the cross-sectional area 198 of thesecondary stage air jets 194 may be at least about 1.75:1, such as atleast about 2:1, or at least about 2.25:1. It should be appreciated,that for embodiments where a cross-section of the primary stage air jets192 and/or secondary stage air jets 194 vary, e.g., along thecircumferential direction C (see, e.g., FIG. 9), the cross-sectionalarea 196 and cross-sectional area 198 refer to an averagecross-sectional area.

In addition to differing in size, the exemplary combustor assembly 100depicted includes a greater number of secondary stage air jets 194 ascompared to the primary stage air jets 192. More particularly, for theembodiment depicted, the combustor assembly 100 defines a ratio of anumber of secondary stage air jets 194 to a number of primary stage airjets 192 of at least about 1.5:1. For example, in certain embodiments,the combustor assembly 100 may define a ratio of a number of secondarystage air jets 194 to a number of primary stage air jets 192 of at leastabout 1.75:1, of at least about 2:1, or of at least about 2.25:1.

Further, at least for the embodiment depicted, the number of primarystage air jets 192 correlates to a number of fuel nozzles 135. Forexample, in the embodiment depicted, the combustor assembly 100 definesa ratio of a number of primary stage air jets 192 to a number of fuelnozzles 135 of at least about 1.5:1. More particularly, in certainembodiments, the combustor assembly 100 may define a ratio of a numberof primary stage air jets 192 to a number of fuel nozzles 135 of atleast about 1.75:1, of at least about 2:1, or of at least about 2.25:1.

Further still, as briefly mentioned above, the quench air jets 176 ofthe inner liner 102 and the outer liner 108 are positioned at theconvergence of the combustion chamber 116, about halfway along a lengthof the inner liner 102 and outer liner 108. For the embodiment depicted,the plurality of primary stage air jets 192 are positioned at a forwardend of the convergence, and the plurality of secondary stage air jets194 are positioned at an aft end of the convergence. Additionally, theprimary stage air jets 192 and secondary stage air jets 194 may beseparated a distance (i.e., a distance from an edge of one opening to anedge of the other opening along the centerline 118) of less than aboutten percent (10%) of a total length of the inner liner 102 or outerliner 108 along the centerline 118. Moreover, as may be seen mostclearly in FIG. 4, the plurality of primary stage air jets 192 areoriented substantially perpendicularly to the centerline 118 of thecombustion chamber 116 (i.e., a length-wise centerline 200 of eachprimary air jet 194 is substantially perpendicular to a local centerline118 of the combustion chamber 116). By contrast, the plurality ofsecondary stage air jets 194 are, for the embodiment, oriented obliquerelative to the centerline 118 of the combustion chamber 116 (i.e., alength-wise centerline 202 of each of the plurality of secondary stageair jets 194 define an angle relative to a local centerline 118 of thecombustion chamber 116). For example, the centerline 118 of each of theplurality of secondary stage air jets 194 may define an angle relativeto a local centerline 118 of the combustion chamber 116 less than aboutseventy-five (75) degrees, such as less than about sixty (60) degrees.It should be appreciated, however, that in other exemplary embodiments,the plurality of primary air jets may instead be oriented oblique to thecenterline 118 of the combustion chamber 116, and further, in certainembodiments, the plurality of secondary stage air jets 194 may beoriented perpendicularly to the centerline 118 of the combustion chamber116.

Notably, aft of the quench air jets 176, the exemplary inner liner 102and outer liner 108 depicted each include a plurality of film coolingholes 206. The film cooling holes 206 provide for cooling of an insidesurface (i.e., a surface adjacent to the combustion chamber 116) of theinner liner 102 and outer liner 108. It should be appreciated, that thefilm cooling holes 206 may be at least about 1/10 of a size of thesecondary stage air jets 194.

Referring now briefly to FIG. 6, providing a close-up, cross-sectionalview of the exemplary primary stage air jet 192 depicted in FIG. 4, theouter liner 108 includes an inlet transition 208 immediately forward ofeach of the plurality of primary stage air jets 192. The inlettransition 208 may reduce an amount of separation of the compressed airfor combustion 134A from the outer liner 108 as the compressed air forcombustion 134A flows over the outer liner 108 and into the primarystage air jets 192. For the embodiment depicted, the inlet transition208 defines a radius of curvature at least about 0.65 inches. Forexample, in certain embodiments, the inlet transition 208 may define aradius of curvature of at least about 0.75 inches. As used herein, theterm “radius of curvature” refers to a radius of a circle that touches acurve at a given point and has the same tangent and curvature at thatpoint.

Further, referring now briefly to FIG. 7, providing an outer view of theprimary stage air jet 192 of the outer liner 108 depicted in FIG. 4, theexemplary primary stage air jet 192 defines an inlet 210 having agenerally elliptical shape. For the embodiment depicted, the ellipticalshape of the inlet includes a minor radius of curvature 212 of at leastabout 0.25 inches and a major radius of curvature 214 of at least about0.4 inches. However, in other embodiments, the elliptical shape of theinlet 210 may instead define any other suitable major and/or minorradius of curvature 212, 214. For example, in other embodiments, theelliptical shape of the inlet 210 may define a minor radius of curvature212 of at least about 0.3 inches and a major radius of curvature 214 ofat least about 0.5 inches. Additionally, it should be appreciated thatin other embodiments, the inlet of the primary stage air jets 192 mayhave any other suitable shape, and similarly, a cross-section of theprimary stage air jets 192 may have any other suitable shape. Forexample, although for the embodiment depicted, the cross-sections of theprimary stage air jets 182 are substantially elliptical, in otherembodiments they may be substantially circular, squared, triangular,polygonal, etc.

By contrast, referring now back to FIG. 5, for the embodiment depicted,each of the plurality of secondary stage air jets 194 are configured aselongated slots, elongated along a widthwise direction 216. Furthermore,for the embodiment depicted, the elongated slots define an oblique angle218 relative to the axial direction A and relative to the centerline 118of the combustion chamber 116 (see particularly, the close up ofelongated slot in FIG. 5). For example, the elongated slots (i.e., thewidthwise direction 216) may define an angle 218 of at least aboutthirty degrees (30°) relative to the axial direction A and centerline118, such as at least about forty-five degrees (45°) relative to theaxial direction A and centerline 118. However, in other embodiments, thewidthwise direction 216 of the elongated slots may define any othersuitable angle 218 with the axial direction A and centerline 118.Additionally, in still other embodiments, it should be appreciated, thatthe secondary stage air jets 194 may instead be circular in shape,elliptical in shape, or have any other suitable shape.

As stated, in certain embodiments, at least about sixty percent (60%) ofthe compressed air for combustion 134A is introduced to the combustionchamber 116 through the quench air jets 176 as a quench airflow 134B.Inclusion of the plurality of primary stage air jets 192 and theplurality secondary stage air jets 194 in accordance with the presentdisclosure may allow for substantially even distribution of the quenchairflow 134B between the plurality of primary stage air jets 192 and theplurality of secondary stage air jets 194. For example, in certainembodiments, the combustor assembly 100 may be configured to providebetween about forty percent (40%) and about sixty percent (60%) of thequench airflow 134B through the plurality of primary stage air jets 192and between about forty percent (40%) and about sixty percent (60%) ofthe quench airflow 134B through the plurality of secondary stage airjets 194. More specifically, in certain embodiments, the combustorassembly 100 may be configured to provide about fifty-five percent (55%)of the quench airflow 134B through the plurality of primary stage airjets 192 and about forty-five (45%) of the quench airflow 134B throughthe plurality of secondary stage air jets 194.

The plurality of quench air jets 176 configured in such a manner mayallow for the quench airflow 134B to reach the combustion gasses fromthe primary combustion zone 172 across an entire height of the secondarycombustion zone 174. More particularly, with such a configuration, theprimary stage air jets 192 may be configured to provide deep penetrationof the quench airflow 134B, while the secondary stage air jets 194 maybe configured to provide relatively shallow penetration of the quenchairflow 134B.

It should be appreciated, that although the description above withreference to FIGS. 4 through 6 is directed to the outer liner 108 andthe quench air jets 176 of the outer liner 108, in certain embodiments,the quench air jets 176 of the inner liner 102 may be configured insubstantially the same manner.

Reference will now be made to FIG. 8. FIG. 8 provides a view along acenterline 118 of the combustion chamber 116 depicted in FIG. 4, lookingforward from an aft end. More particularly, FIG. 8 provides a view ofthe outer liner 108, along with the primary stage air jets 192 andsecondary stage air jets 194, the combustion chamber 116, and an aft endof two of the plurality of fuel nozzles 135. As is depicted, and as maybe seen in other of the Figures, the plurality of primary stage air jets192 are each spaced along the circumferential direction C, andsimilarly, the plurality of secondary stage air jets 194 are also spacedalong the circumferential direction C. Particularly for the embodimentdepicted, each of the plurality of primary stage air jets 192 and theplurality of secondary stage air jets 194 are substantially evenlyspaced along the circumferential direction C. As may also be seen inFIG. 8, each of the plurality of primary stage air jets 192 aresubstantially the same size (i.e., each defines substantially the samecross-sectional area) and further, each of the plurality of secondarystage air jets 194 are also substantially the same size (i.e., eachdefines substantially the same cross-sectional area).

It should be appreciated, however, that in other exemplary embodimentsof the present disclosure, one or both of the plurality of primary stageair jets 192 or the plurality of secondary stage air jets 194 may definea variable spacing along the circumferential direction C and/or avariable size. For example, referring now to FIG. 9, a view is providedalong a centerline 118 of a combustion chamber 116 of a combustorassembly 100, looking forward from an aft end, in accordance withanother exemplary embodiment of the present disclosure is provided.

For the embodiment of FIG. 9, at least one of the plurality of primarystage air jets 192 or plurality of secondary stage air jets 194 areunevenly spaced along the circumferential direction C. Moreparticularly, for the embodiment of FIG. 9, the plurality of secondarystage air jets 194 are unevenly spaced along the circumferentialdirection C. For the embodiment depicted, the uneven spacing along thecircumferential direction C of the plurality of secondary stage air jets194 correlates to a position of the plurality of fuel nozzles 135.Specifically, for the embodiment depicted, the plurality of secondarystage air jets 194 are unevenly spaced along the circumferentialdirection C such that they are more closely spaced immediatelydownstream of the respective fuel nozzles 135.

Moreover, for the embodiment depicted, at least one of the pluralityprimary stage air jets 192 or plurality of secondary stage air jets 194include a variable size along the circumferential direction C, which forthe embodiment depicted also correlates to a position of the pluralityof fuel nozzles 135. Specifically, for the embodiment depicted, theplurality primary stage air jets 192 include a variable sizing along thecircumferential direction C, such that the primary stage air jets 192located immediately downstream of a respective fuel nozzle 135 arelarger (i.e. define a larger cross-sectional area) as compared to otherprimary stage air jets 192 not located immediately downstream of arespective fuel nozzle 135.

It should be appreciated, however, that in other exemplary embodiments,the combustor assembly 100 may have any other suitable configuration,and that the above description is not meant to be limiting unless theclaims specifically provide for such limitations.

A combustor assembly 100 in accordance with one or more embodiments ofthe present disclosure may allow for a more efficient combustor assembly100, with reduced emissions. More particularly, a combustor assembly 100in accordance with one or more embodiments of the present disclosure mayallow for reduced NOx emissions. For example, by including a primarycombustion zone 172 with a relatively high equivalence ratio, andimmediately downstream including a secondary combustion zone 174 with arelatively low equivalence ratio (facilitated by the two-stage quenchair jets 176, lack of forward cooling holes, and/or pre-mixing nozzles),the relatively high combustion temperatures which may generate a maximumamount of NOx may be minimized.

Further, inclusion of the two-stage quench air jets 176 may allow for aplurality primary stage air jets 192 to achieve relatively deeppenetration of a quench airflow into the combustion chamber 116, whilethe plurality of secondary stage air jets 194 may provide for relativelyshallow penetration of a quench airflow into the combustion chamber 116,such that a more uniform provision of quench airflow to the secondarycombustion zone 174 may be provided

Referring now to FIG. 10, a flow diagram is provided of an exemplarymethod (300) for operating a combustor assembly of a gas turbine engine.In certain exemplary aspects, the method (300) may operate one or moreof the exemplary combustor assemblies described above with reference toFIGS. 2 through 8. Accordingly, the exemplary combustor assembly mayinclude a liner attached to or formed integrally with a dome, with theliner including a plurality of quench air jets and the dome including afuel nozzle attached thereto.

The exemplary method (300) includes at (302) providing a substantiallyhomogenous mixture of fuel and air having an equivalence ratio of atleast about 1.5 through the fuel nozzle to a combustion chamber definedat least in part by the dome and the liner. Providing the substantiallyhomogenous mixture of fuel and air at (302) may include providing amixture of fuel and air to the combustion chamber having a mixedness ofat least about seventy percent (70%), such as at least about eightypercent (80%).

Additionally, the exemplary method (300) includes at (304) providing thecombustion chamber with a quench airflow through the plurality of quenchair jets of the liner at a location downstream from the fuel nozzle. Forthe exemplary aspect depicted, the exemplary method is operable with acombustor assembly wherein the liner is an outer liner and wherein thecombustor assembly further includes an inner liner. The outer liner andinner liner each include a plurality of quench air jets positionedbetween a respective forward end and aft end of the liners, with eachset of quench air jets including a plurality of primary stage air jetsspaced from a plurality of secondary stage air jets along an axialdirection and along a centerline of the combustion chamber. Accordingly,for the exemplary aspect depicted, providing the combustion chamber witha quench airflow at (304) includes at (306) providing between aboutforty percent (40%) and about sixty percent (60%) of the quench airflowthrough the plurality of primary stage air jets and at (308) providingbetween about forty percent (40%) and about sixty percent (60%) of thequench airflow through the plurality of secondary stage air jets.

A combustor assembly operated in accordance with exemplary methoddescribed herein may allow for more efficient operation of the combustorassembly with reduced emissions.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A rich-quench-lean combustor assembly for a gasturbine engine defining an axial direction and a circumferentialdirection, the combustor assembly comprising: a fuel nozzle; a dome, thefuel nozzle attached to the dome; and a liner attached to or formedintegrally with the dome, the liner and the dome together defining atleast in part a combustion chamber, the liner extending between a firstforward end and a first aft end, the liner comprising a first annularforward section disposed adjacent the first forward end, a first annularaft section disposed adjacent the first aft end and extending parallelto the first annular forward section, and a first annular convergencesection interconnecting the first annular forward section and the firstannular aft section, the first annular convergence section disposed atan oblique angle to the first annular forward section and the firstannular aft section, the liner comprising a plurality of first quenchair jets positioned between the first forward end and first aft end, thefirst quench air jets comprising a plurality of first primary stage airjets and a plurality of first secondary stage air jets, the plurality offirst primary stage air jets each spaced from the plurality of firstsecondary stage air jets along the axial direction; and wherein theplurality of first primary stage air jets are positioned at a forwardportion of the convergence section, and are oriented substantiallyperpendicularly to the axial direction, and the plurality of firstsecondary stage air jets are positioned at an aft portion of the firstconvergence section, and are oriented oblique relative to the axialdirection such that air is injected in an upstream direction relative toa flow through the combustion chamber, wherein at least one of theplurality of first primary stage air jets defines a firstcross-sectional area, and at least one of the plurality of firstsecondary stage air jets defines a second cross-sectional area less thanthe first cross-sectional area.
 2. The combustor assembly of claim 1,wherein the liner is an outer liner, wherein the combustor assemblyfurther comprises: an inner liner attached to or formed integrally withthe dome, the inner liner defining at least in part the combustionchamber, the inner liner extending between a second forward end and asecond aft end, the inner liner comprising a second annular forwardsection disposed adjacent the second forward end, a second annular aftsection disposed adjacent the second aft end and extending parallel tothe second annular forward section, and a second annular convergencesection interconnecting the second annular forward section and thesecond annular aft section, the second annular convergence sectiondisposed at an oblique angle to the second annular forward section andthe second annular aft section, the inner liner comprising a pluralityof second quench air jets positioned between the second forward end andthe second aft end, the second quench air jets comprising a plurality ofsecond primary stage air jets and a plurality of second secondary stageair jets, the plurality of second primary stage air jets each spacedfrom the plurality of second secondary stage air jets along the axialdirection, wherein the plurality of second primary stage air jets arepositioned at a forward portion of the second convergence section, andare oriented substantially perpendicularly to the axial direction, andthe plurality of second secondary stage air jets are positioned at anaft portion of the second convergence section, and are oriented obliquerelative to the axial direction, wherein at least one of the pluralityof second primary stage air jets defines a third cross-sectional area,and at least one of the plurality of second secondary stage air jetsdefines a fourth cross-sectional area less than the thirdcross-sectional area, wherein the combustor assembly is configured toreceive compressed air for combustion, wherein at least about sixtypercent (60%) of the compressed air for combustion is introduced intothe combustion chamber through the first and second quench air jets as aquench airflow.
 3. The combustor assembly of claim 2, wherein thecombustor assembly is configured to provide between about forty percent(40%) and about sixty percent (60%) of the quench airflow through thefirst and second primary stage air jets and between about forty percent(40%) and about sixty percent (60%) of the quench airflow through thefirst and second secondary stage air jets.
 4. The combustor assembly ofclaim 1, wherein the combustor assembly defines a ratio of a number ofthe first secondary stage air jets to a number of the first primarystage air jets of at least about 2:1.
 5. The combustor assembly of claim1, wherein the plurality of first primary stage air jets are each spacedalong the circumferential direction, and wherein the plurality of firstsecondary stage air jets are each spaced along the circumferentialdirection.
 6. The combustor assembly of claim 5, wherein each of theplurality of first primary stage air jets and the plurality of firstsecondary stage air jets are evenly spaced along the circumferentialdirection.
 7. The combustor assembly of claim 5, wherein at least one ofthe plurality of first primary stage air jets or the plurality of firstsecondary stage air jets are unevenly spaced along the circumferentialdirection.
 8. The combustor assembly of claim 7, wherein the fuel nozzlecomprises a plurality of fuel nozzles spaced evenly along thecircumferential direction, and wherein the uneven spacing along thecircumferential direction of the at least one of the plurality of firstprimary stage air jets or the plurality of first secondary stage airjets correlates to a position of the plurality of fuel nozzles.
 9. Thecombustor assembly of claim 7, wherein the fuel nozzle comprises aplurality of fuel nozzles spaced evenly along the circumferentialdirection, and wherein at least one of the plurality of first primarystage air jets or the plurality of first secondary stage air jets varyin size along the circumferential direction, the variation in sizecorrelating to a position of the plurality of fuel nozzles.
 10. Thecombustor assembly of claim 1, wherein the combustion chamber defines acenterline, wherein the plurality of first secondary stage air jets areeach configured as elongated slots having a widthwise direction definingan oblique angle relative to the centerline.
 11. The combustor assemblyof claim 1, wherein the liner includes an inlet transition immediatelyforward of each of the plurality of first primary stage air jets,wherein the inlet transition defines a radius of curvature of at leastabout 0.65 inches.
 12. The combustor assembly of claim 1, wherein eachof the plurality of first primary stage air jets defines an inlet havingan elliptical shape, wherein the elliptical shape of the inlet includesa minor radius of curvature of at least about 0.25 inches and a majorradius of curvature of at least about 0.4 inches.
 13. The combustorassembly of claim 1, wherein the liner is an outer liner, wherein thecombustor assembly further comprises: an inner liner attached to orformed integrally with the dome, the inner liner defining at least inpart the combustion chamber, the inner liner extending between a secondforward end and a second aft end, the inner liner comprising a secondannular forward section disposed adjacent the second forward end, asecond annular aft section disposed adjacent the second aft end andextending parallel to the second annular forward section, and a secondannular convergence section interconnecting the second annular forwardsection and the second annular aft section, the second annularconvergence section disposed at an oblique angle to the second annularforward section and the second annular aft section, the inner linercomprising a plurality of second quench air jets positioned between thesecond forward end and the second aft end, the second quench air jetscomprising a plurality of second primary stage air jets and a pluralityof second secondary stage air jets, the plurality of second primarystage air jets each spaced from the plurality of second secondary stageair jets along the axial direction, wherein the plurality of secondprimary stage air jets are positioned at a forward portion of the secondconvergence section, and are oriented substantially perpendicularly tothe axial direction, and the plurality of second secondary stage airjets are positioned at an aft portion of the second convergence section,and are oriented oblique relative to the axial direction, wherein atleast one of the plurality of second primary stage air jets defines athird cross-sectional area, and at least one of the plurality of secondsecondary stage air jets defines a fourth cross-sectional area less thanthe third cross-sectional area.
 14. The combustor assembly of claim 13,wherein the combustor assembly defines a forward height within thecombustion chamber between the outer liner and the inner liner at alocation forward of the plurality of first and second quench air jets,wherein the combustor assembly defines an aft height within thecombustion chamber between the outer liner and the inner liner at alocation aft of the plurality of first and second quench air jets, andwherein a ratio of the forward height to the aft height is at leastabout 1.75:1.
 15. The combustor assembly of claim 1, wherein the domeand the liner are each formed of a ceramic matrix composite material.16. A gas turbine engine defining an axial direction and acircumferential direction, the gas turbine engine comprising: acompressor section and a turbine section arranged in serial flow order;and a rich-quench-lean combustor assembly positioned between thecompressor section and the turbine section, the combustor assemblycomprising: a fuel nozzle; a dome, the fuel nozzle attached to the dome;and a liner attached to or formed integrally with the dome, the linerand the dome together defining at least in part a combustion chamber,the liner extending between a first forward end and a first aft end, theliner comprising a first annular forward section disposed adjacent thefirst forward end, a first annular aft section disposed adjacent thefirst aft end and extending parallel to the first annular forwardsection, and a first annular convergence section interconnecting thefirst annular forward section and the first annular aft section, thefirst annular convergence section disposed at an oblique angle to thefirst annular forward section and the first annular aft section, theliner comprising a plurality of first quench air jets positioned betweenthe first forward end and first aft end, the first quench air jetscomprising a plurality of first primary stage air jets and a pluralityof first secondary stage air jets, the plurality of first primary stageair jets each spaced from the plurality of first secondary stage airjets along the axial direction; and wherein the plurality of firstprimary stage air jets are positioned at a forward portion of theconvergence section, and are oriented substantially perpendicularly tothe axial direction, and the plurality of first secondary stage air jetsare positioned at an aft portion of the first convergence section, andare oriented oblique relative to the axial direction such that air isinjected in an upstream direction relative to a flow through thecombustion chamber, wherein at least one of the plurality of firstprimary stage air jets defines a first cross-sectional area, and atleast one of the plurality of first secondary stage air jets defines asecond cross-sectional area less than the first cross-sectional area.17. The gas turbine engine of claim 16, wherein the fuel nozzle isconfigured to provide a mixture of fuel and air to the combustionchamber having at least about a seventy percent (70%) mixedness.
 18. Thegas turbine engine of claim 16, wherein the liner is an outer liner,wherein the combustor assembly further comprises: an inner linerattached to or formed integrally with the dome, the inner liner definingat least in part the combustion chamber, the inner liner extendingbetween a second forward end and a second aft end, the inner linercomprising a second annular forward section disposed adjacent the secondforward end, a second annular aft section disposed adjacent the secondaft end and extending parallel to the second annular forward section,and a second annular convergence section interconnecting the secondannular forward section and the second annular aft section, the secondannular convergence section disposed at an oblique angle to the secondannular forward section and the second annular aft section, the innerliner comprising a plurality of second quench air jets positionedbetween the second forward end and the second aft end, the second quenchair jets comprising a plurality of second primary stage air jets and aplurality of second secondary stage air jets, the plurality of secondprimary stage air jets each spaced from the plurality of secondsecondary stage air jets along the axial direction, wherein theplurality of second primary stage air jets are positioned at a forwardportion of the second convergence section, and are orientedsubstantially perpendicularly to the axial direction, and the pluralityof second secondary stage air jets are positioned at an aft portion ofthe second convergence section, and are oriented oblique relative to theaxial direction, wherein at least one of the plurality of second primarystage air jets defines a third cross-sectional area, and at least one ofthe plurality of second secondary stage air jets defines a fourthcross-sectional area less than the third cross-sectional area, whereinthe combustor assembly is configured to receive compressed air forcombustion, wherein at least about sixty percent (60%) of the compressedair for combustion is introduced into the combustion chamber through thefirst and second quench air jets as a quench airflow.
 19. The gasturbine engine of claim 18, wherein the combustor assembly is configuredto provide between about forty percent (40%) and about sixty percent(60%) of the quench airflow through the first and second primary stageair jets and between about forty percent (40%) and about sixty percent(60%) of the quench airflow through the first and second secondary stageair jets.